1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to leading edge cooling of airfoils in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section in which a high temperature gas flow passes through the rotor blades and stationary vanes or nozzles to extract energy from the flow. The efficiency of the gas turbine engine can be increased by providing a higher flow temperature into the turbine. However, the temperature is limited to the material capabilities of the parts exposed to the hot gas flow.
One method of allowing for higher turbine temperatures is to provide cooling of the first and second stages of the turbine. Complex internal cooling passages have been disclosed to provide high cooling capabilities while using lower flow volumes of cooling air. Since the cooling air used in the turbine airfoils is typically bleed off air from the compressor, using less bleed off air for cooling also increases the efficiency of the engine.
A Prior Art first stage turbine blade is shown in FIG. 1. The turbine blade 10 includes a cooling air supply cavity 11 along the leading edge with drilled film cooling holes forming a leading edge showerhead 12 arrangement and suction side film cooling holes 13 supplied with cooling air form the supply cavity 11. a mid-chord cooling supply channel 14 supplies cooling air to a 3-pass serpentine flow cooling circuit with a second leg 15 and a third leg 16 in which each of the three channels 14,15,16 includes pressure side film cooling holes 17 to discharge film cooling air from the respective channel onto the airfoil surface to provide film cooling. A trailing edge cooling supply channel 18 supplies cooling air to the trailing edge and discharges cooling air through pressure side film cooling holes and trailing edge exit holes 20 arranged along the trailing edge of the airfoil. Exit cooling slots could also be used to discharge the cooling air from the supply channel 18 and out the trailing edge region of the airfoil. FIG. 2 shows a cross section side view of the prior art turbine blade of FIG. 1 with the three cooling supply channels and the 3-pass serpentine flow cooling circuit in the mid-chord region of the blade. Cooling air is also discharged out the blade tip through blade tip cooling holes as shown by the arrows in FIG. 2.
In the prior art first stage turbine blade leading edge cooling construction of FIG. 1 and FIG. 2, a single pass radial flow cooling circuit is used for the airfoil leading edge region. However, the single pass radial flow cooling channel with the drilled film cooling holes design is not the best method of utilizing the cooling air and results in a low convective cooling effectiveness.
U.S. Pat. No. 7,011,502 B2 issued to Lee et al on Mar. 14, 2006 entitled THERMAL SHIELD TURBINE AIRFOIL. Discloses an airfoil with a longitudinal first inlet channel (56 in this patent) connected by a row of impingement holes (48 in this patent) to a longitudinal channel (42 in this patent), being connected by a plurality of film cooling holes (50 in this patent) to the leading edge surface of the airfoil. In the Lee et al patent, the longitudinal channel is also connected to bridge channels on the pressure side and suction sides of the longitudinal channel. Only one multi-impingement channel is used in the Lee et al invention to supply film cooling holes as opposed to the three separate multi-impingement channels of the present invention.
U.S. Pat. No. 4,859,147 issued to Hall et al on Aug. 22, 1989 entitled COOLED GAS TURBINE BLADE discloses an airfoil with a showerhead arrangement having a cooling air supply cavity (24 in this patent), an impingement cavity (28 in this patent) connected to the cooling supply cavity by three slots (26 in this patent), and three film cooling holes (30 in this patent) connected to the impingement cavity.
U.S. Pat. No. 6,379,118 B2 issued to Lutum et al on Apr. 30, 2002 entitled COOLED BLADE FOR A GAS TURBINE discloses a similar showerhead arrangement to that above of the Hall et al patent. A cooling air supply cavity (50 in this patent) is connected to an impingement cavity (47 in this patent) through two impingement cooling holes (49 in this patent), and three film cooling holes (48 in this patent) are connected to the impingement cavity. Both of the Hall et al and Lutum et al patents lack the first impingement cavity and the second impingement cavity in series, and the multi-impingement cavities of the present invention.